Staggered core printout

ABSTRACT

A core for gas turbine engine component comprises a body extending between first and second ends to define a length, and extending between first and second edges to define a width. A plurality of core extensions are formed as part of the body. The plurality of core extensions are positioned to be staggered relative to each other such that at least two adjacent core extensions are variable relative to each other in at least one dimension. A gas turbine engine component is also disclosed.

BACKGROUND OF THE INVENTION

In pursuit of higher engine efficiencies, higher turbine inlettemperatures have been relied upon to boost overall engine performance.This can result in gas path temperatures that may exceed melting pointsof turbine component constituent materials. To address this issue,dedicated cooling air is extracted from a compressor section and is usedto cool the gas path components in the turbine, such as rotating bladesand stator vanes for example, incurring significant cycle penalties.

One method of cooling turbine airfoils utilizes internal coolingchannels or cavities formed in the airfoil to promote convective heattransfer. Cooling air is typically routed from a root of the airfoiltoward a tip. The cooling air is then discharged out of the airfoilthrough a plurality of holes formed along a length of the airfoil. Thecooling air exiting the holes forms a film of cooler air that shieldsthe airfoil from incoming combustion gases.

Typically, these cooling holes and significant cooling mass flow ratesare required to provide the needed amount of cooling. In order toeffectively cool the airfoils to protect against damage, there is a needto balance the amount of cooling flow used and the overall heat transfercapability.

SUMMARY OF THE INVENTION

In a featured embodiment, a core for a gas turbine engine componentcomprises a body extending between first and second ends to define alength, and extending between first and second edges to define a width.A plurality of core extensions are formed as part of the body. Theplurality of core extensions are positioned to be staggered relative toeach other such that at least two adjacent core extensions are variablerelative to each other in at least one dimension.

In another embodiment according to the previous embodiment, the bodycomprises a curvilinear structure.

In another embodiment according to any of the previous embodiments, thecore extensions are spaced apart from each other.

In another embodiment according to any of the previous embodiments, apoint along one of the first and second edges defines a radius ofcurvature. The core extensions are positioned on the radius ofcurvature.

In another embodiment according to any of the previous embodiments, thecore extensions are spaced apart from each other long the radius ofcurvature.

In another embodiment according to any of the previous embodiments, atleast one dimension comprises a length of the extrusion in a generalradial direction, and wherein at least two of the core extensions havelengths that are different from each other.

In another embodiment according to any of the previous embodiments, thecore extensions comprise polygonal or curved shaped core extensions.

In another embodiment according to any of the previous embodiments, thebody defines an airfoil internal passage shape.

In another featured embodiment, a gas turbine engine component comprisesa body extending between first and second ends to define a length andextending between first and second edges to define a width. A pluralityof openings are formed within a wall surface of the body. The pluralityof openings are positioned to be staggered relative to each other suchthat at least two adjacent openings are offset from each other in atleast one direction.

In another embodiment according to any of the previous embodiments, apoint along one of the first and second edges defines a radius ofcurvature. The openings are positioned on the radius of curvature andare spaced apart from each other.

In another embodiment according to any of the previous embodiments, thebody comprises an airfoil with the first edge comprising a leading edge,the second edge comprising a trailing edge, the first end comprising aradially inner portion and the second end comprising a tip portion. Thepoint is located at a position that is nearer to the leading edge thanthe trailing edge.

In another embodiment according to any of the previous embodiments, flowexiting the openings is generally parallel or angled relative to anassociated streamline path.

In another embodiment according to any of the previous embodiments, theopenings comprise polygonal or curved openings.

In another embodiment according to any of the previous embodiments, theopenings are formed in a pressure and/or suction side of the airfoil.

In another embodiment according to any of the previous embodiments, theopenings are formed in a platform for a vane, blade, or BOAS.

In another embodiment according to any of the previous embodiments, theopenings are spaced apart from each other along an arcuate pathextending between the first and second ends.

In another embodiment according to any of the previous embodiments, theopenings comprise polygonal or curved openings.

In another embodiment according to any of the previous embodiments, thebody comprises one of an airfoil, a blade, a vane, a BOAS, or acombustor panel.

In another embodiment according to any of the previous embodiments, thebody comprises an airfoil with the first edge comprising a leading edge,the second edge comprising a trailing edge, the first end comprising aradially inner portion and the second end comprising a tip portion. Thepoint is located at a position that is nearer to the trailing edge thanthe leading edge.

In another embodiment according to any of the previous embodiments, thebody comprises an airfoil with the first edge comprising a leading edge,the second edge comprising a trailing edge, the first end comprising aradially inner portion and the second end comprising a tip portion, andwherein the point is located at a position that is nearer to thepressure side than the suction side.

In another embodiment according to any of the previous embodiments, thebody comprises an airfoil with the first edge comprising a leading edge,the second edge comprising a trailing edge, the first end comprising aradially inner portion and the second end comprising a tip portion. Thepoint is located at a position that is nearer to the suction side thanthe pressure side.

In another embodiment according to any of the previous embodiments, bodycomprises an airfoil with the first edge comprising a leading edge, thesecond edge comprising a trailing edge, the first end comprising aradially inner portion and the second end comprising a tip portion. Thepoint is located at a position that nearer to the engine centreline thana radial outward location.

In another featured embodiment, a method of manufacturing a gas turbineengine component includes providing a body extending between first andsecond ends to define a length and extending between first and secondedges to define a width. A plurality of openings is formed within a wallsurface of the body. The plurality of openings are positioned to bestaggered relative to each other such that at least two adjacentopenings are offset from each other in at least one direction. Flowexiting the openings is generally parallel or angled relative to anassociated streamline path, and forms the openings via one of a casting,EDM, laser, or additive manufacturing method.

The foregoing features and elements may be combined in any combinationwithout exclusivity, unless expressly indicated otherwise.

These and other features may be best understood from the followingdrawings and specification.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a schematic representation of one example of a gas turbineengine.

FIG. 2 is a side perspective view of a turbine blade.

FIG. 3 is a schematic cross-sectional view of a root section of theblade of FIG. 2.

FIG. 4 is a schematic view of a core printout design for an airfoil.

FIG. 5A is a schematic perspective view of a core with a staggered printout as shown in FIG. 4.

FIG. 5B shows the core as used in an airfoil section.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbofan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. Alternative engines mightinclude an augmentor section (not shown) among other systems orfeatures. The fan section 22 drives air along a bypass flow path B in abypass duct defined within a nacelle 15, while the compressor section 24drives air along a core flow path C for compression and communicationinto the combustor section 26 then expansion through the turbine section28. Although depicted as a two-spool turbofan gas turbine engine in thedisclosed non-limiting embodiment, it should be understood that theconcepts described herein are not limited to use with two-spoolturbofans as the teachings may be applied to other types of turbineengines including three-spool architectures.

The exemplary engine 20 generally includes a low speed spool 30 and ahigh speed spool 32 mounted for rotation about an engine centrallongitudinal axis A relative to an engine static structure 36 viaseveral bearing systems 38. It should be understood that various bearingsystems 38 at various locations may alternatively or additionally beprovided, and the location of bearing systems 38 may be varied asappropriate to the application.

The low speed spool 30 generally includes an inner shaft 40 thatinterconnects a fan 42, a first (or low) pressure compressor 44 and afirst (or low) pressure turbine 46. The inner shaft 40 is connected tothe fan 42 through a speed change mechanism, which in exemplary gasturbine engine 20 is illustrated as a geared architecture 48 to drivethe fan 42 at a lower speed than the low speed spool 30. The high speedspool 32 includes an outer shaft 50 that interconnects a second (orhigh) pressure compressor 52 and a second (or high) pressure turbine 54.A combustor 56 is arranged in exemplary gas turbine 20 between the highpressure compressor 52 and the high pressure turbine 54. A mid-turbineframe 57 of the engine static structure 36 is arranged generally betweenthe high pressure turbine 54 and the low pressure turbine 46. Themid-turbine frame 57 further supports bearing systems 38 in the turbinesection 28. The inner shaft 40 and the outer shaft 50 are concentric androtate via bearing systems 38 about the engine central longitudinal axisA which is collinear with their longitudinal axes.

The core airflow is compressed by the low pressure compressor 44 thenthe high pressure compressor 52, mixed and burned with fuel in thecombustor 56, then expanded over the high pressure turbine 54 and lowpressure turbine 46. The mid-turbine frame 57 includes airfoils 59 whichare in the core airflow path C. The turbines 46, 54 rotationally drivethe respective low speed spool 30 and high speed spool 32 in response tothe expansion. It will be appreciated that each of the positions of thefan section 22, compressor section 24, combustor section 26, turbinesection 28, and fan drive gear system 48 may be varied. For example,gear system 48 may be located aft of combustor section 26 or even aft ofturbine section 28, and fan section 22 may be positioned forward or aftof the location of gear system 48.

The engine 20 in one example is a high-bypass geared aircraft engine. Ina further example, the engine 20 bypass ratio is greater than about six(6), with an example embodiment being greater than about ten (10), thegeared architecture 48 is an epicyclic gear train, such as a planetarygear system or other gear system, with a gear reduction ratio of greaterthan about 2.3 and the low pressure turbine 46 has a pressure ratio thatis greater than about five. In one disclosed embodiment, the engine 20bypass ratio is greater than about ten (10:1), the fan diameter issignificantly larger than that of the low pressure compressor 44, andthe low pressure turbine 46 has a pressure ratio that is greater thanabout five 5:1. Low pressure turbine 46 pressure ratio is pressuremeasured prior to inlet of low pressure turbine 46 as related to thepressure at the outlet of the low pressure turbine 46 prior to anexhaust nozzle. The geared architecture 48 may be an epicycle geartrain, such as a planetary gear system or other gear system, with a gearreduction ratio of greater than about 2.3:1. It should be understood,however, that the above parameters are only exemplary of one embodimentof a geared architecture engine and that the present invention isapplicable to other gas turbine engines including direct driveturbofans.

A significant amount of thrust is provided by the bypass flow B due tothe high bypass ratio. The fan section 22 of the engine 20 is designedfor a particular flight condition—typically cruise at about 0.8 Mach andabout 35,000 feet. The flight condition of 0.8 Mach and 35,000 ft, withthe engine at its best fuel consumption—also known as “bucket cruiseThrust Specific Fuel Consumption (‘TSFC’)”—is the industry standardparameter of lbm of fuel being burned divided by lbf of thrust theengine produces at that minimum point. “Low fan pressure ratio” is thepressure ratio across the fan blade alone, without a Fan Exit Guide Vane(“FEGV”) system. The low fan pressure ratio as disclosed hereinaccording to one non-limiting embodiment is less than about 1.45. “Lowcorrected fan tip speed” is the actual fan tip speed in ft/sec dividedby an industry standard temperature correction of [(Tram ° R)/(518.7°R)]^(0.5). The “Low corrected fan tip speed” as disclosed hereinaccording to one non-limiting embodiment is less than about 1150ft/second.

Airfoils located downstream of combustor section 26, such as statorvanes and rotor blades in the turbine section 28 for example, operate ina high-temperature environment. Airfoils that are exposed to hightemperatures typically include internal cooling channels that direct aflow of cooling air through the airfoil to remove heat and prolong theuseful life of the airfoil. FIG. 2 is a side view of a turbine rotorblade 60 having a root section 62, a platform 64, and an airfoil section66. Root section 62 is connected to a rotor in the turbine section 28(FIG. 1) as known. The airfoil section 66 includes a leading edge 68, atrailing edge 70, a suction side wall 72, and a pressure side wall 74.The airfoil section 66 extends to a tip 76, and includes a plurality ofsurface cooling holes, such as film cooling holes 78, and a plurality oftrailing edge cooling slots 79.

The platform 64 connects one end of airfoil section 66 to root section62. The leading edge 68, trailing edge 70, suction side wall 72, andpressure side wall 74 extend outwardly away from the platform 64. Thetip 76 closes off an opposite end of the airfoil section 66 from theplatform 64. Suction side wall 72 and pressure side wall 74 connectleading edge 68 and trailing edge 70. Film cooling holes 78 are arrangedover a surface of airfoil section 66 to provide a layer of cool airproximate the surface of airfoil section 66 for protection fromhigh-temperature combustion gases. Trailing edge cooling slots 79 arearranged along trailing edge 70 to provide an exit for air circulatingwithin airfoil section 66.

FIG. 3 is a schematic section view of the root section 62 of the rotorblade 60 of FIG. 2. The rotor blade 60 includes one or more internalcooling channels. In the example shown, there is at least a firstcooling channel 81 near the leading edge 68, and a second coolingchannel 83 positioned aft of the first cooling channel 81. The coolingchannels 81, 83 direct cooling flow F radially outwardly toward the tip76 of the blade 60. The cooling channels 81, 83 deliver cooling flow tothe film cooling holes 78 and the cooling slots 79. The cooling channelsinternal to the airfoil section 66 can take various forms.

FIGS. 4 and 5A-5B show a structure 100 for providing a sacrificial corethat is used in making gas turbine engine components such as airfoilsfor a blade or vane, a blade outer air seal (BOAS), or a combustorpanel, for example. As known, the core is used to define a shapedopening within the finished component. For example, the core is used todefine the internal cooling channels 81, 83 the airfoil section 66.Typically, the core is formed from a ceramic material or refractorymetal; however, other suitable materials could also be used.

In order to improve cooling efficiency, the subject invention provides acore containing a feature that will leave a staggered core printout.This witness will follow streamline patterns such that the printouts areparallel or angled with the streamline, providing optimum film decay andaero mixing losses. This will result in an improvement in airfoil life.

As shown in FIGS. 5A-5B, a core 80 is used to form internal passageswithin a gas turbine engine component, such as an airfoil section 66.The shape of the core 80 is a positive structure that forms acorresponding negative shaped feature within the airfoil section 66. Thecore 80 comprises a body 82 that extends between a first end 84 and asecond end 86 to define a length L, and extends between a first edge 88and a second edge 90 to define a width W. A plurality of core extensions92 is formed within as part of the body 82. The core extensions 92 arepositioned to be staggered relative to each other such that adjacentcore extensions 92 are variable relative to each other in a directionextending from the first end 84 to the second end 86 and or in adirection extending from the first edge 88 to a second edge 90.

In other words, at least two adjacent core extensions 92 are configuredsuch that they are variable from each other in at least one dimension.For example, as shown in FIG. 5A, lengths of the core extensions 92along the radial direction are different from each other. Longer coreextensions are located near the center while shorter core extensions arelocated at the ends 84, 86. This is merely one example configuration,and other configurations could also be utilized.

The length L of the body 80 may or may not be greater than the width W.In one example, ends of the core extensions 92 are spaced apart fromeach other along a radius of curvature R. The spacing is defined by apoint P that is located along one of the first 88 and second 90 edges.The core extensions 92 are positioned on the radius R. As such, the coreextensions 92 are circumferentially spaced apart from each other long anarc segment S.

In one example, the body 82 is used to form an internal passage for anairfoil. In this example, the point P is positioned at the first edge 88which corresponds to the leading edge 68 of the airfoil section 66.Further, as shown, the body 82 comprises a curvilinear structureextending between the edges 88, 90 to form a passage that includes bothcurved and/or or straight portions (FIG. 5B).

In one example, the core extensions 92 comprise rectangular-shaped coreextensions that are defined by a first dimension extending along the arcsegment and a second dimension extending transverse to the arc segmentS. The first dimension, which corresponds generally to a length of theprotrusion 92, is greater than the second dimension, which correspondsgenerally to a width of the protrusion. While rectangular shaped coreextensions are shown as an example, it should be understood that othershapes could also be used such as polygonal, oval, or round shapes, forexample.

The body 82 comprises a sacrificial component that defines an airfoilinternal passage shape. As discussed above, the airfoil section 66provides a body that extends from a radially inner end 100 to a radiallyouter end 102 to define a length, and extends between leading 68 andtrailing 70 edges to define a width. The core extensions 92 containedwithin the core body 82 form a corresponding plurality of openings 104within a wall surface 106 of the airfoil section 66. The openings 104are open to the external surface and are staggered relative to eachother as shown in FIG. 4. In one example, adjacent openings 104 are notaligned with each other in a direction extending from the radially innerend 100 to the radially outer end 102; however, other configurationscould also be used.

As discussed above, the position of the openings 104 is determined bythe location of the point P, which defines a radius of curvature R. Theopenings 104 are positioned along the arc segment S defined by theradius of curvature R and are circumferentially spaced apart from eachother.

As known, streamline paths are defined in a direction generallyextending from the leading edge 68 to the trailing edge 70 and possiblyextending outward toward the suction side 72. The position and shapes ofthe openings 104 allows the cooling air flow exiting the openings 104 tobe generally parallel or angled relative to an associated streamlinepath. This provides an optimized amount of film cooling.

In the example shown, the openings 104 are formed in the pressure side74 of the airfoil section 66. Openings 104 could also be formed in thesuction side 72 or openings 110 (FIG. 2) could be formed in the platform64. Further, while the core body 82 is disclosed as forming an airfoilsection 66 for a blade, the core body could also be used to form a vane,a blade outer air seal (BOAS), or a combustor panel for example.Additionally, instead of using cores, casting, EDM, laser, or additivemanufacturing methods could be used to form the openings 104.

Although an embodiment of this invention has been disclosed, a worker ofordinary skill in this art would recognize that certain modificationswould come within the scope of this invention. For that reason, thefollowing claims should be studied to determine the true scope andcontent of this invention.

The invention claimed is:
 1. A gas turbine engine component comprising:a body extending between first and second ends to define a length andextending between first and second edges to define a width; a pluralityof openings formed within a wall surface of the body, wherein theplurality of openings are positioned to be staggered relative to eachother such that at least two adjacent openings are offset from eachother in at least one direction; and an arc segment formed along thelength of the body, wherein the openings are spaced apart from eachother along the arc segment.
 2. The component according to claim 1including a point along one of the first and second edges that defines aradius of curvature for the arc segment.
 3. The component according toclaim 2 wherein the body comprises an airfoil with the first edgecomprising a leading edge, the second edge comprising a trailing edge,the first end comprising a radially inner portion and the second endcomprising a tip portion, and wherein the point is located at a positionthat is nearer to the leading edge than the trailing edge.
 4. Thecomponent according to claim 3 wherein flow exiting the openings isgenerally parallel or angled relative to an associated streamline path.5. The component according to claim 4 wherein the body includes aninternal cooling channel and a passage that leads from the internalcooling channel to an external surface of the body, and wherein theopenings comprise polygonal or curved openings in the external surface.6. The component according to claim 4 wherein the openings are formed ina pressure and/or suction side of the airfoil.
 7. The componentaccording to claim 4 wherein the openings are formed in a platform for avane, blade, or blade outer air seal.
 8. The component according toclaim 1 wherein the openings are spaced apart from each other along thearc segment extending between the first and second ends.
 9. Thecomponent according to claim 8 wherein the body includes an internalcooling channel and a passage that leads from the internal coolingchannel to an external surface of the body, and wherein the openingscomprise polygonal or curved openings in the external surface.
 10. Thecomponent according to claim 1 wherein the body comprises one of anairfoil, a blade, a vane, a blade outer air seal, or a combustor panel.11. The component according to claim 2 wherein the body comprises anairfoil with the first edge comprising a leading edge, the second edgecomprising a trailing edge, the first end comprising a radially innerportion and the second end comprising a tip portion, and wherein thepoint is located at a position that is nearer to the trailing edge thanthe leading edge.
 12. The component according to claim 2 wherein thebody comprises an airfoil with the first edge comprising a leading edge,the second edge comprising a trailing edge, the first end comprising aradially inner portion and the second end comprising a tip portion, andwherein the point is located at a position that is nearer to a pressureside than a suction side.
 13. The component according to claim 2 whereinthe body comprises an airfoil with the first edge comprising a leadingedge, the second edge comprising a trailing edge, the first endcomprising a radially inner portion and the second end comprising a tipportion, and wherein the point is located at a position that is nearerto a suction side than a pressure side.
 14. The component according toclaim 2 wherein the body comprises an airfoil with the first edgecomprising a leading edge, the second edge comprising a trailing edge,the first end comprising a radially inner portion and the second endcomprising a tip portion, and wherein the point is located at a positionthat nearer to an engine centerline than the tip portion.
 15. Thecomponent according to claim 1, wherein the arc segment curves betweenfirst and second arc ends, and wherein openings at each of the first andsecond arc ends are closer to one of the first and second edges thanremaining openings that are located between the openings at each of thefirst and second arc ends.
 16. The component according to claim 1,wherein the body comprises an airfoil with the first edge comprising aleading edge, the second edge comprising a trailing edge, the first endcomprising a radially inner portion and the second end comprising a tipportion, and including a point along the leading edge that defines aradius of curvature for the arc segment such that openings at a centerportion of the arc segment are nearer to the trailing edge than theopenings at each opposing end of the arc segment.